Although this invention is applicable to the incorporation of various electronic devices in the structure of a vehicle, it has been found to be particularly useful in the environment of the incorporation of antennas and/or sensors by embedding the antennas and/or sensors in aircraft structure and mounting related electronics/optics, as appropriate, to the backplane. Therefore, without limiting the applicability of the invention to "the incorporation of antennas by embedding the antennas in the aircraft structure", the invention will be described in such environment. The detailed description which follows is for the example of an antenna structure incorporated in an aircraft
Typically, today's aircraft antenna systems are not designed to carry structural loads. These structurally parasitic prior art antenna systems adversely affect the overall airframe weight with a corresponding reduction in aircraft performance and fuel consumption. The placement of antennas in aircraft structure is presently limited to locations which are lightly loaded. Basically an opening is cut in the aircraft for mounting an antenna therein. At that location, the load-bearing capability of the structure is reduced. To return the load-bearing capability to the structure, the structure around the opening is normally increased in thickness and overall weight so the load may be carried around the opening. This extra weight, relative to the normal weight of the undamaged or uncut area of the aircraft skin, is undesirable.
It is common practice in the aircraft industry to use a sandwich-type or composite-type structure in the fabrication of aircraft wherein a primary load-bearing skin is provided on the outside, then a core of honeycomb-type or similar material and then a second skin on the inside of the core to provide a stiffer material which is usually stronger and/or stiffer on a weight basis. Although this type of structure provides greater stiffness, the disadvantage with non-metallic designs is that damage to the structure may not be readily apparent and therefore the structure must be over-designed with a damage tolerance criteria. Typically, this type of structure has a large margin of safety included in the design in order to provide for the unknown factor of damage caused by such factors as a worker dropping a wrench on the structure, stones and debris being kicked up from the runway and hitting the structure and the like.
If damage to the outer load-bearing skin of the composite-type structure of the aircraft is visible and is noticed, repair of that damage may require a lengthy and complicated procedure. The damaged structure will normally be moved to a repair facility having the high temperature and high pressure apparatus necessary to make a repair which will ensure that the aircraft has the same structural integrity after the repair as the aircraft had before the damage.
Due to placement limitations imposed by structural load-carrying requirements, antennas and other electronics and avionics are adversely restricted to less than optimum performance. In most cases, antenna o locations are limited to lightly loaded locations where a wingtip location or the leading edge of the horizontal or vertical stabilizer might be a much better location from the standpoint of antenna gain, radio frequency coverage, sensor coverage and the like. When prior art antennas and other electronics and avionics are located at wing-tip or leading edge locations for better overall performance, then the airframe is adversely affected by additional structure and weight.
It is also well known to install an antenna in an aircraft and then surround the antenna with a radome to provide aerodynamic flow around that antenna. The radome is not considered to be part of the primary load path and does not contribute to the load-bearing capability of the aircraft structure. The radome will transfer local airloads to the connecting structure.
The present invention is intended to provide a solution to various prior art deficiencies which include antennas and antenna systems which do not contribute to supporting the structural load in the aircraft.